orbitalElements

Orbital elements of satellites in scenario

Syntax

elements = orbitalElements(sat)

Description

elements = orbitalElements(sat) returns the orbital elements of the specified satellite sat.

Examples

Retrieve Orbital Elements of Satellite

Create a satellite scenario object.

sc = satelliteScenario;

Add a satellite to the scenario.

tleFile = "eccentricOrbitSatellite.tle";
sat1 = satellite(sc,tleFile);

Retrieve the orbital elements of sat1.

elements1 = orbitalElements(sat1)

elements1 = struct with fields:
                       MeanMotion: 1.4544e-04
                     Eccentricity: 0.7415
                      Inclination: 60.0000
    RightAscensionOfAscendingNode: 30.0000
              ArgumentOfPeriapsis: 280
                      MeanAnomaly: 289.4697
                           Period: 43200
                            Epoch: 05-May-2020 13:51:55
                            BStar: 0

Add a satellite from Keplerian elements to the scenario.

semiMajorAxis = 6878137;                               % meters
eccentricity = 0;
inclination = 20;                                      % degrees
rightAscensionOfAscendingNode = 0;                     % degrees
argumentOfPeriapsis = 0;                               % degrees
trueAnomaly = 0;                                       % degrees
sat2 = satellite(sc,semiMajorAxis,eccentricity, ...
       inclination,rightAscensionOfAscendingNode, ...
       argumentOfPeriapsis,trueAnomaly, ...
       "OrbitPropagator","two-body-keplerian", ...
       "Name","Sat2");

Retrieve the orbital elements of sat2.

elements2 = orbitalElements(sat2)

elements2 = struct with fields:
                    SemiMajorAxis: 6878137
                     Eccentricity: 0
                      Inclination: 20
    RightAscensionOfAscendingNode: 0
              ArgumentOfPeriapsis: 0
                      TrueAnomaly: 0
                           Period: 5.6770e+03

Input Arguments

sat — Satellite

row vector of Satellite objects

Satellite, specified as a row vector of Satellite objects.

Output Arguments

elements — Orbital elements

structure

Orbital elements of the input sat, returned as a structure. The fields of the structure depend on the orbit propagator you specify using the OrbitPropagator property of the satelliteScenario object.

Two-Body Keplerian

The two-body-keplerian orbit propagator returns these fields.

• SemiMajorAxis, in meters

• Eccentricity

• Inclination, in degrees

• RightAscensionOfAscendingNode, in degrees

• ArgumentOfPeriapsis, in degrees

• TrueAnomaly, in degrees

• Period, in seconds

SGP4 and SDP4

The sgp4 and sdp4 orbit propagators returns these fields.

• MeanMotion, in degrees/second

• Eccentricity

• Inclination, in degrees

• RightAscensionOfAscendingNode, in degrees

• ArgumentOfPeriapsis, in degrees

• MeanAnomaly, in degrees

• Epoch

• BStar, in 1/EarthRadius

• Period, in seconds

The orbital elements represent the mean values at Epoch.

Ephemeris

The ephemeris propagator returns these fields.

• EphemerisStartTime

• EphemerisStopTime

• PositionTimeTable

• VelocityTimeTable

GPS

The gps propagator returns these fields.

• PRN

• SatelliteHealth

• GPSWeekNumber

• GPSTimeOfApplicability, in seconds

• SemiMajorAxis, in meters

• Eccentricity

• Inclination, in degrees

• GeographicLongitudeOfOrbitalPlane, in degrees

• RateOfRightAscension, in degrees/second

• ArgumentOfPerigee, in degrees

• MeanAnomaly, in degrees

• Period, in seconds

The orbital elements are derived from the SEM almanac file and defined in the Earth-Centered-Earth-Fixed (ECEF) frame.

Version History

Introduced in R2021a